1. Field of the Invention
High-thrust, liquid propellant rocket engines capable of delivering a payload to orbit.
2. Brief Description of the Prior Art
The design of liquid propellant rocket engines capable of delivering a payload to orbit is an extremely complex and difficult art. Chemical (liquid) propellants possess limited energy. It is difficult to design an engine that utilizes this energy effectively enough to deliver a payload to orbit. The most effective known rocket engines provide thrust-to-weight ratios that are only slightly greater than those provided by engines incapable of delivering any payload to orbit.
Rocket engine design is further complicated by the fact that each component of a rocket engine interacts with every other component in a very complex manner. Attempts to improve one component or portion of an engine often produce an undesirable effect in some other portion of the engine that more than offsets the attempted improvement.
One relatively effective known rocket engine capable of delivering a payload to orbit comprises a thrust chamber, fuel feed and oxidizer feed pipelines, a single turbopump assembly for pumping both fuel and oxidizer through those pipelines to the thrust chamber, and a single preburner combustion chamber for generating a gaseous discharge that drives the turbopump. The preburner combustion chamber is disposed in the fuel supply pipeline and connected to receive a small quantity of oxidizer to provide controlled fuel-rich combustion. In operation, the preburner generates sufficient heat to convert the liquid fuel to a gas. Gaseous discharge from the preburner drives the turbopump which in turn pumps both fuel and oxidizer to the combustion chamber. Heat exchanger apparatus transfers heat from the preburner to the oxidizer feed pipeline so that both fuel and oxidizer are converted to gases before entering the thrust chamber. The gaseous fuel and oxidizer mix rapidly in the thrust chamber and interact combustibly to generate a gaseous output flow that produces a high thrust. Embodiments of this engine burning a hydrocarbon fuel and using liquid oxygen oxidizer are known that inject propellant into the thrust chamber at up to 1,300 psia and provide a thrust-to-weight ratio as high as approximately 90 at sea level.
The problems presented to one attempting to design a new rocket engine or improve an existing one can be appreciated by considering some of the many complex interrelationships between the different components of the rocket engine described above. The thrust generated by that engine is determined at least in part by the pressure at which the propellants are injected into the thrust chamber. Thrust can be increased by increasing the pressures at which the propellants are supplied to the thrust chamber. Propellant injection pressure is dependent upon the operating temperature of the preburner combustion chamber and can be increased by raising the operating temperature of that chamber. But, if the operating temperature of the preburner combustion chamber is raised to too high a level in an attempt to increase thrust, additional apparatus will be required for cooling the walls of the preburner combustion chamber. This additional cooling apparatus may be so heavy that it more than offsets the desired increased thrust. And, increasing the pressure at which propellants are injected into the thrust chamber may also require alteration of other portions of the engine. For example, in the above-described engine, the oxidizer is directed through cooling passageways in the thrust chamber wall before being injected into the thrust chamber. The oxidizer thus absorbs heat and thereby cools the thrust chamber wall. Changes in the injection pressure of any propellant also change the mass flow and therefore the heat transfer characteristics of that propellant so that the thrust chamber cooling scheme may have to be modified. Changes in propellant injection pressure may also require a larger, heavier turbopump to accommodate the higher temperature and pressure gaseous discharge from the preburner and also to provide the desired higher injection pressures. Thus, a straightforward attempt to increase the thrust-to-weight ratio of an engine by increasing the operating temperature of a preburner to thereby increase the pressures at which the fuel and oxidizer are injected into the thrust chamber may require changes in other engine components such that the thrust-to-weight ratio of the engine is actually reduced. Any engine modification altering the pressure at which propellant is injected into the thrust chamber must affect the other components of the engine in a manner that produces an overall beneficial result.
The complicated interrelationships between different rocket engine components that make it difficult to increase propellant injection pressure also make it difficult to achieve other objectives. Safety, for example, is an important consideration. Since one turbopump assembly is used to pump both propellants, very heavy and sophisticated seals are needed to prevent any leakage across that turbine that could mix combustibly interactive propellants and produce an explosion. These seals increase the weight and complexity of the engine. But, turbopump modifications cannot be made considering only the safety factor. Engine weight, thrust, propellant injection pressure, propellant mass flow, and heat transfer factors must also be considered.
All prior art liquid propellant rocket engines use only one fuel and one oxidizer to generate thrust. It has been recognized that an engine capable of utilizing different propellants having different densities during different portions of a rocket flight would be substantially superior to an engine that burns only one fuel. Such an engine would make a single stage vehicle for delivering a payload to earth orbit and then returning, a practical reality. A number of individuals have pointed out the advantages and desirability of developing such an engine. But, no one has been able to design a practical embodiment. The problems in obtaining a desirable balance between the structural and operational characteristics of the different interrelated components of a rocket engine, as described above for an engine using one fuel and one oxidizer, are compounded significantly when additional propellants having different densities and thus different masses, momentums, flow characteristics, and heat transfer characteristics are considered.
The prior art does not teach how an effective propellant delivery system for pumping different propellants having different densities, mass flow, heat transfer, and other characteristics, to a thrust chamber should be designed. That is, the prior art does not indicate whether the propellant delivery system in an engine for effectively utilizing different propellants during different portions of a rocket flight should be similar to one of the many different propellant delivery systems used in different engines designed to operate with only one fuel and one oxidizer and, if so, which one, or whether it should be different from all of those prior art systems.
The prior art does not teach an effective scheme for injecting different propellants into the thrust chamber during different portions of a flight. The injection orifices on the face of an injector in an engine that burns a high-density fuel are spaced differently from those on an injector in an engine for burning a low-density fuel. Significant problems must be overcome with any injector design approach that one might select in attempting to provide a rocket engine capable of using different propellants. For example, if one designing an engine attempts to maintain the difference between injector spacing in engines that use high density and low density propellants and inject the different propellants into the same thrust chamber through different orifices, he risks a heat distribution across the injector face that will burn out the injector during any portion of a rocket flight when one propellant is either not being used or is being used only in small quantities. But, on the other hand, if he attempts to inject different propellants having different densities through the same orifices during a different portion of a rocket flight, he faces difficult problems in insuring that the different propellant combinations having different densities, mass flow rates, momentum ratios, and so forth, each mix in a desirable manner and provide stable, high thrust generating combustion. The prior art does not teach how to accomplish this without encountering an unacceptable weight or performance penalty in some portion of the engine.
The prior art also does not teach how an engine designed to utilize different propellant combinations during different portions of a rocket flight should be cooled. There are significant advantages and drawbacks to any cooling scheme that can be hypothesized. Consider, for example, an engine designed to use hydrocarbon fuel, which is a dense fuel, for one portion of a flight, and hydrogen, which is a low density fuel, for another. Liquid oxygen is an oxidizer for both fuels. Hydrogen has a much higher heat capacity than hydrocarbon fuel, and oxygen, and therefore would be the best coolant. But, the drawback in attempting to use hydrogen as the coolant is that it is not used to provide thrust during all portions of a flight. A cooling system designed to use different propellants for cooling during different portions of a flight could be considered. The drawback involved is that it would have to be relatively complicated. Oxygen is supplied to the engine whenever either one of the fuels are, and therefore is available for cooling. The drawback involved is that it has a much lower heat capacity than hydrogen. The prior art does not teach what cooling structure should be used to achieve both good cooling and also operate in good harmony with the other components of a rocket engine designed to use different propellants during different portions of a mission.